Low noise turbine for geared turbofan engine

ABSTRACT

A gas turbine engine is utilized in combination with a gear reduction to reduce the speed of a fan relative to a low pressure turbine speed. The gas turbine engine is designed such that a blade count in the low pressure turbine multiplied by the speed of the low pressure turbine will result in operational noise that is above a sensitive range for human hearing. A method and turbine module are also disclosed.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.13/403,005, filed Feb. 23, 2012 now U.S. Pat. No. 8,246,292, whichclaims priority to U.S. Provisional Application No. 61/592,643, filedJan. 31, 2012.

BACKGROUND

This application relates to the design of a turbine which can beoperated to produce noise that is less sensitive to human hearing.

Gas turbine engines are known, and typically include a fan deliveringair into a compressor. The air is compressed in the compressor anddelivered downstream into a combustor section where it was mixed withfuel and ignited. Products of this combustion pass downstream overturbine rotors, driving the turbine rotors to rotate.

Typically, there is a high pressure turbine rotor, and a low pressureturbine rotor. Each of the turbine rotors include a number of rows ofturbine blades which rotate with the rotor. Interspersed between therows of turbine blades are vanes.

The low pressure turbine can be a significant noise source, as noise isproduced by fluid dynamic interaction between the blade rows and thevane rows. These interactions produce tones at a blade passage frequencyof each of the low pressure turbine stages, and their harmonics.

The noise can often be in a frequency range that is very sensitive tohumans. To mitigate this problem, in the past, a vane-to-blade ratio hasbeen controlled to be above a certain number. As an example, avane-to-blade ratio may be selected to be 1.5 or greater, to prevent afundamental blade passage tone from propagating to the far field. Thisis known as “cut-off.”

However, acoustically cut-off designs may come at the expense ofincreased weight and reduced aerodynamic efficiency. Stated another way,by limiting the designer to a particular vane to blade ratio, thedesigner may be restricted from selecting such a ratio based upon othercharacteristics of the intended engine.

Historically, the low pressure turbine has driven both a low pressurecompressor section and a fan section. More recently, a gear reductionhas been provided such that the fan and low pressure compressor can bedriven at distinct speeds.

SUMMARY OF THE INVENTION

In a featured embodiment, a gas turbine engine has a fan, a compressorsection including a low pressure compressor and a high pressurecompressor, a combustor section, a high pressure turbine and a lowpressure turbine. The low pressure turbine drives the low pressurecompressor and fan. A gear reduction is utilized to reduce the speed ofthe fan from an input speed from the low pressure turbine. The lowpressure turbine has a number of turbine blades in each of a pluralityof rows of the turbine. The low pressure turbine blades operate at arotational speed. The number of blades and the rotational speed areselected such that the following formula holds true for at least one ofthe blade rows of the low pressure turbine: (number ofblades×speed)/60≧5500. The rotational speed is an approach speed inrevolutions per minute.

In another embodiment according to the foregoing embodiment, the formularesults in a number greater than or equal to 6000.

In another embodiment according to the foregoing embodiment, the gasturbine engine is rated to produce 15,000 pounds of thrust or more.

In another embodiment according to the immediately foregoing embodiment,the formula holds true for the majority of the blade rows of the lowpressure turbine.

In another embodiment according to the immediately foregoing embodiment,the formula holds true for all of the blade rows of the low pressureturbine.

In another embodiment according to the featured embodiment, the formulaholds true for the majority of the blade rows of the low pressureturbine.

In another embodiment according to the featured embodiment, the formulaholds true for all of the blade rows of the low pressure turbine.

In another featured embodiment, a method of designing a gas turbineengine includes the steps of including a gear reduction between a lowpressure turbine and a fan, and selecting the number of blades of thelow pressure turbine rotors, in combination with the rotational speed ofthe low pressure turbine, such that the following formula holds true forat least one of the blade rows of the low pressure turbine: (number ofblades×speed)/60≧5500. The rotational speed is an approach speed inrevolutions per minute.

In another embodiment according to the foregoing embodiment, the formularesults in a number greater than or equal to 6000.

In another embodiment according to the foregoing embodiment, the gasturbine engine is rated to produce 15,000 pounds of thrust or more.

In another embodiment according to the immediately foregoing embodiment,the formula holds true for the majority of the blade rows of the lowpressure turbine.

In another embodiment according to the immediately foregoing embodiment,the formula holds true for all of the blade rows of the low pressureturbine.

In another embodiment according to the featured embodiment, the formulaholds true for the majority of the blade rows of the low pressureturbine.

In another embodiment according to the immediately foregoing embodiment,the formula holds true for all of the blade rows of the low pressureturbine.

In another featured embodiment, a turbine module for a gas turbineengine has a low pressure turbine with a number of turbine blades ineach of a plurality of rows of the turbine. The low pressure turbineblades operate at a rotational speed. The number of blades and therotational speed are selected such that the following formula holds truefor at least one of the blade rows of the low pressure turbine: (numberof blades×speed)/60≧5500. The rotational speed is an approach speed inrevolutions per minute.

In another embodiment according to the foregoing embodiment, the formularesults in a number greater than or equal to 6000.

In another embodiment according to the foregoing embodiment, the gasturbine engine is rated to produce 15,000 pounds of thrust or more.

In another embodiment according to the immediately foregoing embodiment,the formula holds true for the majority of the blade rows of the lowpressure turbine.

In another embodiment according to the immediately foregoing embodiment,the formula holds true for all of the blade rows of the low pressureturbine.

In another embodiment according to the featured embodiment, the formulaholds true for the majority of the blade rows of the low pressureturbine.

In another embodiment according to the featured embodiment, the formulaholds true for all of the blade rows of the low pressure turbine. Theseand other features of the invention would be better understood from thefollowing specifications and drawings, the following of which is a briefdescription.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a gas turbine engine.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown), or an intermediate spool,among other systems or features. The fan section 22 drives air along abypass flowpath while the compressor section 24 drives air along a coreflowpath for compression and communication into the combustor section 26then expansion through the turbine section 28. Although depicted as aturbofan gas turbine engine in the disclosed non-limiting embodiment, itshould be understood that the concepts described herein are not limitedto use with turbofans as the teachings may be applied to other types ofturbine engines including three-spool architectures.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 50 thatinterconnects a high pressure compressor 52 and high pressure turbine54. A combustor 56 is arranged between the high pressure compressor 52and the high pressure turbine 54. A mid-turbine frame 57 of the enginestatic structure 36 is arranged generally between the high pressureturbine 54 and the low pressure turbine 46. The mid-turbine frame 57further supports bearing systems 38 in the turbine section 28. The innershaft 40 and the outer shaft 50 are concentric and rotate via bearingsystems 38 about the engine central longitudinal axis A which iscollinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

The terms “low” and “high” as applied to speed or pressure for thespools, compressors and turbines are of course relative to each other.That is, the low speed spool operates at a lower speed than the highspeed spool, and the low pressure sections operate at lower pressurethan the high pressures sections.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than ten (10), the gearedarchitecture 48 is an epicyclic gear train, such as a planetary gearsystem or other gear system, with a gear reduction ratio of greater thanabout 2.3 and the low pressure turbine 46 has a pressure ratio that isgreater than about 5. In one disclosed embodiment, the engine 20 bypassratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout 5:1. Low pressure turbine 46 pressure ratio is pressure measuredprior to inlet of low pressure turbine 46 as related to the pressure atthe outlet of the low pressure turbine 46 prior to an exhaust nozzle.The geared architecture 48 may be an epicycle gear train, such as aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.5:1. It should be understood, however, that theabove parameters are only exemplary of one embodiment of a gearedarchitecture engine and that the present invention is applicable toother gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a Fan Exit Guide Vane(“FEGV”) system. The low fan pressure ratio as disclosed hereinaccording to one non-limiting embodiment is less than about 1.45. “Lowcorrected fan tip speed” is the actual fan tip speed in ft/sec dividedby an industry standard temperature correction of [(Tambient degR)/518.7)^0.5]. The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

The use of the gear reduction between the low pressure turbine spool andthe fan allows an increase of speed to the low pressure compressor. Inthe past, the speed of the low pressure turbine has been somewhatlimited in that the fan speed cannot be unduly large. The maximum fanspeed is at its outer tip, and in larger engines, the fan diameter ismuch larger than it may be in smaller power engines. However, the use ofthe gear reduction has freed the designer from limitation on the lowpressure turbine speed caused by a desire to not have unduly high fanspeeds.

It has been discovered that a careful design between the number ofrotating blades, and the rotational speed of the low pressure turbinecan be selected to result in noise frequencies that are less sensitiveto human hearing.

A formula has been developed as follows:(blade count×rotational speed)/60≧5500 Hz.

That is, the number of rotating blades in any low pressure turbinestage, multiplied by the rotational speed of the low pressure turbine(in revolutions per minute), divided by 60 (to put the amount persecond, or Hertz) should be greater than or equal to 5500 Hz. Morenarrowly, the amount should be above 6000 Hz.

The operational speed of the low pressure turbine as utilized in theformula should correspond to the engine operating conditions at eachnoise certification point defined in Part 36 or the FederalAirworthiness Regulations. More particularly, the rotational speed maybe taken as an approach certification point as defined in Part 36 of theFederal Airworthiness Regulations. For purposes of this application andits claims, the term “approach speed” equates to this certificationpoint.

It is envisioned that all of the rows in the low pressure turbine meetthe above formula. However, this application may also extend to lowpressure turbines wherein the majority of the blade rows in the lowpressure turbine meet the above formula, but perhaps some may not.

This will result in operational noise that would be less sensitive tohuman hearing.

In embodiments, it may be that the formula can result in a range ofgreater than or equal to 5500 Hz, and moving higher. Thus, by carefullydesigning the number of blades and controlling the operational speed ofthe low pressure turbine (and a worker of ordinary skill in the artwould recognize how to control this speed) one can assure that the noisefrequencies produced by the low pressure turbine are of less concern tohumans.

This invention is most applicable to jet engines rated to produce 15,000pounds of thrust or more. In this thrust range, prior art jet engineshave typically had frequency ranges of about 4000 hertz. Thus, the noiseproblems as mentioned above have existed.

Lower thrust engines (<15,000 pounds) may have operated under conditionsthat sometimes passed above the 4000 Hz number, and even approached 6000Hz, however, this has not been in combination with the gearedarchitecture, nor in the higher powered engines which have the largerfans, and thus the greater limitations on low pressure turbine speed.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

The invention claimed is:
 1. A gas turbine engine comprising: a fan anda turbine having a fan drive rotor; a gear reduction effecting areduction in the speed of said fan relative to an input speed from saidfan drive rotor; said fan drive rotor having a number of turbine bladesin at least one of a plurality of rows of said fan drive rotor, and saidturbine blades operating at least some of the time at a rotationalspeed, and said number of turbine blades in said at least one row andsaid rotational speed being such that the following formula holds truefor said at least one row of the fan drive turbine(number of blades×speed)/60≧5500 Hz; and said rotational speed being inrevolutions per minute.
 2. The gas turbine engine as set forth in claim1, wherein the formula results in a number greater than or equal to 6000Hz.
 3. The gas turbine engine as set forth in claim 2, wherein said gasturbine engine is rated to produce 15,000 pounds of thrust or more. 4.The gas turbine engine as set forth in claim 1, wherein the formulaholds true for the majority of blade rows of the fan drive rotor.
 5. Thegas turbine engine as set forth in claim 1, wherein said rotationalspeed being an approach speed.
 6. The gas turbine engine as set forth inclaim 1, wherein said turbine section having a higher pressure turbinerotor and a lower pressure turbine rotor, with said fan drive rotorbeing said lower pressure turbine rotor.
 7. A method of designing a gasturbine engine comprising the steps of: including a gear reductionbetween a fan drive turbine rotor and a fan, and selecting a number ofblades in at least one row of the fan drive turbine rotor, incombination with a rotational speed of the fan drive turbine rotor, suchthat the following formula holds true for said at least one row of thefan drive turbine rotor:(number of blades×speed)/60≧5500 Hz; and said rotational speed being inrevolutions per minute.
 8. The method of designing a gas turbine engineas set forth in claim 7, wherein the formula results in a number greaterthan or equal to
 6000. 9. The method of designing a gas turbine engineas set forth in claim 8, wherein said gas turbine engine is rated toproduce 15,000 pounds of thrust or more.
 10. The method as set forth inclaim 7, wherein the formula holds true for the majority of the bladerows of the fan drive turbine.
 11. The method as set forth in claim 7,wherein said rotational speed is an approach speed.
 12. The method asset forth in claim 7, wherein a turbine section including a higherpressure turbine rotor and a lower pressure turbine rotor, and said fandrive turbine rotor being said lower pressure turbine rotor.
 13. Aturbine module comprising: a fan drive rotor having a first blade rowthat includes a number of blades, the first blade row being capable ofrotating at a rotational speed, so that when measuring said rotationalspeed in revolutions per minute:(number of blades×said rotational speed)/60≧5500 Hz.
 14. The turbinemodule as set forth in claim 13, wherein the formula results in a numbergreater than or equal to
 6000. 15. The turbine module as set forth inclaim 13, wherein the formula holds true for the majority of blade rowsof the fan drive rotor.
 16. The turbine module as set forth in claim 13,wherein the rotational speed is an approach speed.
 17. The turbinemodule as set forth in claim 13, wherein there being a higher pressureturbine rotor and a lower pressure turbine rotor, and said fan driverotor being said lower pressure turbine rotor.